Needled ceramic matrix composite cooling passages

ABSTRACT

A method for forming a passage in a ceramic matrix composite component incudes forming a core for a ceramic matrix composite component; embedding a hollow member into the core at a desired location for a passage in the ceramic matrix composite component; wrapping the core with a ceramic material; and inserting a rod through the hollow member and into the core.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a divisional of U.S. patent application Ser. No.15/863,158, filed Jan. 5, 2018.

BACKGROUND

The present disclosure relates to ceramic matrix composite components,and more particularly, to a cooling passage therein.

Gas turbine engines typically include a compressor section to pressurizeairflow, a combustor section to burn a hydrocarbon fuel in the presenceof the pressurized air, and a turbine section to extract energy from theresultant combustion gases. Gas path components, such as turbine blades,often include airfoil cooling that may be accomplished by external filmcooling, internal air impingement, and forced convection, eitherseparately, or in combination.

Ceramic matrix composite (CMC) components can endure high temperatures,but those temperatures may be below the gas path operationaltemperatures of some modern turbine engine stages. Internal convectivecooling of CMC components may be primarily from impingement baffle likestructures and film cooling from cooling passages. The cooling passagesare either drilled using laser or Electron Discharge Machining (EDM).Laser passages are ideal for relatively short passages such as those atthe leading edge of an airfoil, while EDM is ideal for long passagessuch as those at the trailing edge. EDM relies on a current supplied toan electrode which is discharged through a grounded part, however CMCmaterial cannot carry current, and cannot use EDM. Machining coolingpassages in the CMC component may result in cut fibers in the CMCmaterial which may weaken the CMC component or expose a surface toenvironmental attacks.

SUMMARY

A method for forming a passage in a ceramic matrix composite component,according to one disclosed non-limiting embodiment of the presentdisclosure includes forming a core for a ceramic matrix compositecomponent; embedding a hollow member into the core at a desired locationto form a passage in the ceramic matrix composite component; wrappingthe core with a ceramic material; and inserting a rod through the hollowmember and into the core.

A further aspect of the present disclosure includes a plurality offibers through which the hollow member extends but does not cut.

A further aspect of the present disclosure includes separating aplurality of fibers around the hollow member.

A further aspect of the present disclosure includes penetrating theceramic material with the hollow member.

A further aspect of the present disclosure includes that the rod ismanufactured of the same material as the core.

A further aspect of the present disclosure includes that the rod is of adesired cooling passage shape.

A further aspect of the present disclosure includes a sharpened end.

A further aspect of the present disclosure includes that the hollowmember is a needle.

A further aspect of the present disclosure includes that an innersurface of hollow member is sized to receive the rod.

A further aspect of the present disclosure includes that the hollowmember is manufactured of metal.

A further aspect of the present disclosure includes coaxially forming ablind hole in the core for the rod within a blind hole in the core forthe hollow member.

A further aspect of the present disclosure includes gluing the rod intothe blind hole.

A further aspect of the present disclosure includes removing the hollowmember and leaving the rod in the core.

A further aspect of the present disclosure includes removing the coreand the rod from the ceramic material.

A further aspect of the present disclosure includes burning out the coreand the rod from the ceramic material.

A further aspect of the present disclosure includes forming the hollowmember of a nylon.

A further aspect of the present disclosure includes burning out thecore, the hollow member, and the rod from the ceramic material.

A ceramic matrix composite component according to one disclosednon-limiting embodiment of the present disclosure includes a ceramicmaterial in which a cooling passage is cast, wherein the cooling passageis cast through a separation of a plurality of ceramic fibers of theceramic material.

A further aspect of the present disclosure includes that the ceramicmatrix composite component is an airfoil of a gas turbine engine.

A further aspect of the present disclosure includes that the ceramicmaterial forms a wall of the airfoil.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of an example gas turbine enginearchitecture;

FIG. 2 is an exploded view of rotor assembly with a singlerepresentative ceramic matrix composite turbine blade;

FIG. 3 is a cross-sectional illustration of an example ceramic matrixcomposite turbine blade of the gas turbine engine.

FIG. 4 is a cross-sectional illustration of an example ceramic matrixcomposite turbine blade taken along line 4-4 in FIG. 3.

FIG. 5 is a cross-sectional illustration of an example ceramic matrixcomposite turbine blade taken along line 4-4 in FIG. 3 illustratingformation of a passage in the ceramic matrix composite componentaccording to one disclosed non-limiting embodiment.

FIG. 6 illustrates a flow diagram of an example method to form a passagein the ceramic matrix composite component according to one disclosednon-limiting embodiment.

FIG. 7 illustrates a cross-sectional view of a core for a ceramic matrixcomposite component illustrating a step of the method of FIG. 6illustrating drilling a counterbored hole.

FIG. 8 illustrates a cross-sectional view of a core for a ceramic matrixcomposite component illustrating a step of the method of FIG. 6illustrating embedding a hollow member into the core.

FIG. 9 illustrates a cross-sectional view of a core for a ceramic matrixcomposite component illustrating a step of the method of FIG. 6illustrating wrapping with a ceramic material such that the hollowmember penetrates therethrough.

FIG. 10 illustrates a cross-sectional view of a core for a ceramicmatrix composite component illustrating a step of the method of FIG. 6illustrating inserting a rod into the hollow member.

FIG. 11 illustrates a cross-sectional view of a core for a ceramicmatrix composite component illustrating a step of the method of FIG. 6illustrating removing the hollow member.

FIG. 12 illustrates a cross-sectional view of a core for a ceramicmatrix composite component illustrating a step of the method of FIG. 6illustrating the ceramic material closing around the rod.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 as disclosed herein is a two spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26, and a turbine section 28. The fan section 22drives air along a bypass flowpath while the compressor section 24drives air along a core flowpath for compression and communication intothe combustor section 26, then expansion through the turbine section 28.Although depicted as a high bypass gas turbofan engine architecture inthe disclosed non-limiting embodiment, it should be appreciated that theconcepts described herein are not limited only thereto.

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation around an engine central longitudinal axis Arelative to an engine case structure 36 via several bearings 38. The lowspool 30 generally includes an inner shaft 40 that interconnects a fan42, a low pressure compressor (“LPC”) 44 and a low pressure turbine(“LPT”) 46. The inner shaft 40 drives the fan 42 directly or through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspool 30. An exemplary reduction transmission is an epicyclictransmission, namely a planetary or star gear system. The high spool 32includes an outer shaft 50 that interconnects a high pressure compressor(“HPC”) 52 and high pressure turbine (“HPT”) 54. A combustor 56 isarranged between the HPC 52 and the HPT 54.

With reference to FIG. 2, a rotor assembly 60 such as a turbine rotorassembly includes an array of blades 84 (one shown) circumferentiallydisposed around a disk 86. Each blade 84 includes a root 88, a platform90 and an airfoil 92. The blade root 88 is received within a rim 94 ofthe disk 86 and the airfoil 92 extends therefrom.

The platform 90 separates a gas path side inclusive of the airfoil 92and a non-gas path side inclusive of the root 88. The airfoil 92 definesa blade chord between a leading edge 98, which may include variousforward and/or aft sweep configurations, and a trailing edge 100. Afirst sidewall 102 that may be convex to define a suction side, and asecond sidewall 104 that may be concave to define a pressure side arejoined at the leading edge 98 and at the axially spaced trailing edge100. The tip 96 extends between the sidewalls 102, 104 opposite theplatform 90.

With reference to FIG. 3, to resist the high temperature stressenvironment in the gas path of a turbine engine, each blade 84 includesan array of internal passageways 110. The array of internal passageways110 includes a multiple of feed passages 112 through the root 88 thatcommunicates airflow into a multiple of cavities 114 (shownschematically) within the airfoil 92. The cavities 114 distribute thecooling flow through passages 130 in the sidewalls 102, 104, leadingedge 98, and/or the trailing edge 100 (also shown in FIG. 4).Impingement passages 132 (FIG. 4) may also be located though internalwalls 134 between one or more of the array of internal passageways 110.It should be appreciated that various feed architectures, cavities, andpassageway arrangements will benefit herefrom.

With reference to FIG. 5, the example cooled turbine airfoil 84 ismanufactured as a CMC component. Though the CMC may have less strengthrelative to metallic counterparts, CMCs can endure higher materialtemperatures and are significantly lighter. Although a turbine bladewill be used to illustrate the disclosed cooling passage formationmethod, other components will also benefit herefrom.

The example turbine airfoil 84 is generally formed from a core 200 whichmay be formed from multiple portions 200A, 200B which are wrapped with aceramic material 202. The core 200 is later removed such that the curedceramic material 202 forms the airfoil 92 and the array of internalpassageways 110. The core 200 may comprise a material such as carbon.The core 200 is readily cast and/or machined with conventional methodsthen later removed without damage to the ceramic material 202. The core200 may include a multiple of longitudinal grooves 201.

The ceramic material 202 may be an arrangement of ceramic fibers 204.Examples of the ceramic material 202 may include a three-dimensionalweave of the ceramic fibers 204. Alternatively, or in addition, theceramic material 202 may include a two-dimensional weave of the ceramicfibers 204. The ceramic material 202 may include multiple layers oftwo-dimensional weave of the ceramic fibers 204. Alternatively, or inaddition, the ceramic material 202 may include a fiber layup, such as aunidirectional layup. In some examples, each of the ceramic fibers 204may be a bundle and/or a tow of ceramic fibers. The fibers in eachbundle or tow may be braided or otherwise arranged. The ceramic fibers204 may comprise a material that is stable at temperatures above 1000degrees Celsius. Examples of the ceramic fibers 204 may include fibersof alumina, mullite, silicon carbide, silicon, zirconia or carbon.

With reference to FIG. 6, a method 300 for forming the passages 130through, for example, the airfoil sidewall 102, 104 (FIG. 4) in aceramic matrix composite component is illustrated in a schematic blockdiagram form. It should also be appreciated that application is notlimited to aerospace components and various other applications willbenefit herefrom.

Once the core 200 is manufactured (302), a counterbored hole 212/216(FIG. 7) is drilled (304; FIG. 6) at each location in which the passages130 are to be formed. The counterbored hole 212/216 includes a blindhole 212 for a rod 214 within a blind hole 216 for a hollow member 218along a common axis 220 (FIG. 5). A step 222 is formed between the blindholes 212, 216 to form a stop for the hollow member 218. Eachcounterbored hole 212/216 is located and oriented to form the respectivepassages 130.

Next, the hollow member 218 is located in each blind hole 216 (306, FIG.6; FIG. 8). The hollow member 218 has an interior diameter equal to orgreater to the size of the desired passage 130. The outside diameter maybe equal to or slightly smaller than the blind hole 216. The hollowmember 218 may include a sharp end 219 to form a hollow needle and maybe manufactured of a metal alloy, a nylon, or any other rigid materialthat is compatible with the CMC material.

Next, the core 200 is wrapped with the ceramic material 202 using thehollow member 218 to pierce through the ceramic material 202 (308, FIG.6; FIG. 9). The ceramic material 202 comprises the plurality of fibers204 through which the hollow member 218 extends but does not cut. Thehollow members 218 are of a strength to penetrate and separate theplurality of fibers 204.

The core 200 is wrapped with the ceramic material 202 to form a ceramicmatrix composite body that may be the CMC component in which the passage130 is to be formed. Alternatively, the ceramic matrix composite bodymay be a component of the CMC component in which the passage 130 is tobe formed. The ceramic matrix composite body may comprise of, forexample, a silicon carbide ceramic matrix composite. The ceramic matrixcomposite body may have any shape or form, not just the shapeillustrated. Once all the layers of the ceramic material 202 are inplace, one rod 214 is inserted into each hollow member 218 (310, FIG. 6;FIG. 10).

The rod 214 is shaped and sized to form the desired passages 130. Therod 214 may be formed of the same material as the core such as a carbon.The rod 214 may be circular, rectilinear, oval, racetrack, or of othercross-sectional shape. Optionally, each rod 214 may be glued into eachblind hole 212 with a glue 223.

Next, the hollow member 218 is removed (312, FIG. 6; FIG. 11) leavingthe rod 214 in place. The ceramic material 202 then closes (FIG. 12)around the rod 214. The ceramic fibers 204 of the ceramic material 202are not broken in this process, such that it is readily apparent thatthis method was used because any drilling method would result in thecutting of the ceramic fibers 204. Alternatively, the hollow member 218can be manufactured of a material such as a nylon which can be readilyburned out with the core 200 and then the rod 214 is burned out so thatthe hollow member need not be removed. That is, the hollow member 218burns out at a lower temperature than infiltration temps then the rod214 burns out with the core 200.

Next, the ceramic material 202 is cured (314, FIG. 6) per conventionalCMC manufacturing procedures to form the CMC component. Forming thecooled turbine airfoil 84 as the CMC component may include infiltratinga molten metal or alloy into the ceramic material 202. The multiple oflongitudinal grooves 201 (FIG. 7) facilitate the infiltration.

The molten metal or alloy fills the gaps between the ceramic fibers 204and the rods 214. The molten metal or alloy may also react with areactive element source present in the ceramic material 202 to form theceramic matrix material. In some examples, a chemical vapor infiltrationcoating may be applied to the ceramic material 202 prior to the meltinfiltration to stiffen the ceramic fibers 204. Alternatively, or inaddition, forming the CMC component from the ceramic material 202 mayinclude chemical vapor infiltrating the ceramic material 202 instead ofmelt infiltrating.

Finally, the core 200 and rods 214 are removed (316, FIG. 6) via heat,acid, or other method which does not harm the ceramic material 202 perconventional CMC manufacturing procedures. Once the core 200 and rods214 are removed, the passages 130 and the array of internal passageways114 are formed.

The “cast in” passages 130 are readily identifiable, may be of variouscross-sectional shapes, reduce machining time, and facilitate themanufacture of long passages through CMC components such as thosethrough the trailing edge of an airfoil.

The use of the terms “a”, “an”, “the”, and similar references in thecontext of description (especially in the context of the followingclaims) are to be construed to cover both the singular and the plural,unless otherwise indicated herein or specifically contradicted bycontext. The modifier “about” used in connection with a quantity isinclusive of the stated value and has the meaning dictated by thecontext (e.g., it includes the degree of error associated withmeasurement of the particular quantity). All ranges disclosed herein areinclusive of the endpoints, and the endpoints are independentlycombinable with each other.

Although the different non-limiting embodiments have specificillustrated components, the embodiments of this invention are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be appreciated that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be appreciated that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reason,the appended claims should be studied to determine true scope andcontent.

1-17. (canceled)
 18. A ceramic matrix composite component, comprising: aceramic material in which a cooling passage is cast, wherein the coolingpassage is cast through a separation of a plurality of ceramic fibers ofthe ceramic material.
 19. The ceramic matrix composite component asrecited in claim 18, wherein the ceramic matrix composite component isan airfoil of a gas turbine engine.
 20. The ceramic matrix compositecomponent as recited in claim 19, wherein the ceramic material forms awall of the airfoil.